1. Field of the Invention
The present invention relates to gyroscopic reference apparatus for navigable craft, such as an aircraft, and more particularly relates to an inertial measurement unit or module for a strapped down attitude and heading reference system and or navigation system for aircraft embodying a unique arrangement and orientation of the axes of two two-degrees-of-freedom gyroscopic rate sensors, so as to provide in one module a passive gyro system or in two modules a fail operational gyro system. The gyro orientations also greatly simplify the aircraft body rate equations as well as the parity or failure detection and isolation equations thereby minimizing computer requirements.
A single-degree-of-freedom (SDF) gyroscopic rate sensor is one in which a spinning mass is so mounted relative to a vehicle that it can detect vehicle rates about but one axis. A two-degree-of-freedom (TDF) gyroscopic rate sensor is one in which a spinning mass is so mounted relative to a vehicle that it can detect vehicle rates about two orthogonal axes. A typical two-degree-of-freedom gyroscopic rate sensor is disclosed in the present assignee's U.S. Pat. No. 3,529,477 issued to T. R. Quermann and in copending U.S. Application Ser. No. 818,486 entitled "Permanent Magnet Torquer for Free Rotor Flexure Suspended Gyroscopes", filed July 25, 1977 in the names of C. Buckley and J. Kiedrowski.
2. Description of the Prior Art
Vehicle inertial reference systems based on a plurality of SDF or TDF rate sensors strapped down to the vehicle structure for measuring vehicle rates of rotation about its coordinate axes in combination with acceleration and direction sensors and computer means for computing vehicle acceleration, rate and displacement relative to the earth's or some other, coordinate axes are well known in the art and have been extensively described and discussed in the literature. For example, such inertial systems have been used extensively in missile and space vehicles. However, their use in commercial aircraft have not heretofore been economically feasible and the more traditional gimbal isolated gyroscopes and gyro platforms have been employed to provide direct measures of aircraft attitude relative to earth axes. However, strapped down gyro systems are now becoming practical from a weight, reliability, maintainability and cost of ownership standpoint with the advent of small, high accuracy and relatively low cost two-axis rate sensors and small, lightweight yet powerful digital computers for performing the computations required for coordinate transformation, integration, gyro drift or "erection" computations and failure detection isolation and conversion computations.
Skewing the axes of rate sensors in order to reduce the number of gyros required to provide redundancy of the rate measures for fail operational or dual fail operational systems is a generally old technique familiar to those skilled in the art of strapped down systems and also described extensively in the literature. For example, in one prior art dual redundant skewing arrangement, six SDF rate sensors are oriented with their sensitive axes lying on the surfaces of a dodecahedron oriented in a predetermined relationship relative to the aircraft coordinate axes so that each gyro measures a known component of aircraft rate about its coordinate axes. This dodecahedron configuration while satisfying redundancy requirements has been found to be very costly and lacking in accuracy. Another known dual redundant arrangement designed for military aircraft is to distribute six separate SDF solid state rate sensors at various spaced locations in the aircraft and to orient them so that their sensitive axes lie on the surface of a cone with a one-half angle of 77 degrees. In this arrangement each rate sensor is physically large and are independently housed apart from the system electronics and computer in groups of one, two or three per package, the packages dispersed about the aircraft to minimize a total system failure due, for example, to battle damage. The sensor dimensions prohibited them from being packaged in a single housing and their rate sensing capability required the very large cone half angle. Furthermore, the rate measurement matrix is very complex and requires substantial computation time. In short, this single degree of freedom rate sensor/cone configuration is unsuitable for application in commercial aircraft. Still another dual fail operational arrangements have been proposed in the prior art; such as, orienting four TDF rate gyros so that their sensitive axes lie in adjacent planes of a semioctahedron. This octahedron orientation, like the dodecahedron orientation, is not as accurate as the conic configuration. Also, in this system one TDF gyro, two accelerometers and associated computer electronics are mounted in a single LRU (line replaceable unit) and the required four LRU's were in turn mounted in a specially designed aircraft rack in different orientations such as to orient the gyros' input axes as described; a complex mechanical arrangement in the rack interconnected the gyros of each LRU so as to simultaneously precisely orient the gyro units relative to one another and to the aircraft axes. All of these systems are very complex mechanically, electrically and mathematically and are very expensive and not readily adaptable to general aviation or commercial aircraft use.
For commercial airline applications, simplicity of hardware, installation and calibration and maintenance removal in terms of minimum LRU's reliability, and over-all cost effectiveness are key considerations and it is to these that the inertial measurement unit (IMU) of the present invention is primarily directed.